Optimal sun safe attitude for satellite ground tracking

ABSTRACT

A method and apparatus for maneuvering a satellite in orbit to alternately optimize the collection of solar energy and to take sensor data of terrestrial objects is disclosed The longitudinal axis of a large payload package is oriented perpendicular to the orbital plane to minimize the disturbance torque due to gravity gradient, and to allow simple rotation about the axis for attitude change between optimal Sun and optimal ground coverage.

BACKGROUND OF THE INVENTION

1. Field of the Invention

The present invention relates generally to systems and methods formaneuvering satellites in orbit, and in particular, to a system andmethod for maneuvering a satellite to direct a one sensors atterrestrially based locations and for optimal collection of solarenergy.

2. Description of the Related Art

Satellites typically comprise solar panels that are used to collectsolar energy. The solar energy capabilities of such panes are maximizedif the panes face directly into the Sun (that is, a vector normal to theplane of the solar panes is directed at the Sun). While solar panels aretypically capable of rotating along one axis, the satellite itselfusually also needs to be oriented to absorb maximum energy from the Sun.

In most circumstances, this poses no significant problems, as the momentof inertia of the satellite is small enough to permit such satelliteorientations without undue requirements placed upon the satellite'sattitude control subsystems. However, some satellite missions demand theuse of sensor packages that substantially increase the satellite'smoment of inertia about at least one of the satellite's axes. Forexample, sensor arrays used for space-based ground surveillance usingradar can be quite large compared to other satellite structures,substantially increasing the satellite's moment of inertia. Such arraysmay also be fixed to the satellite bus itself, thereby requiting thesatellite bus to rotate to direct the arrays at the desired terrestriallocation. As a result, such satellites can be asked to perform frequentand large rotations about axes with large moments of inertia. Thisplaces costly requirements on the satellite attitude control system,particularly the motive elements (e.g. thrusters, momentum wheels andthe like) that are used to rotate the satellite.

What is needed is a system and method that permits satellites with largeor extended payloads to achieve their mission while still permittingmaximum solar energy absorption without the use of high capacity motiveelements. The present invention satisfies this need.

SUMMARY OF THE INVENTION

To address the requirements described above, the present inventiondiscloses a method and apparatus for maneuvering a satellite in orbit toalternately optimize the collection of solar energy and to take sensordata of terrestrial objects. The satellite has a payload such as asensor disposed longitudinally along a first (x) axis and rotatablesolar panels disposed longitudinally along a second (y) axisperpendicular to the first axis (x), the satellite having a moment ofinertia about the first axis I_(XX) and about the second axis I_(YY) anda moment of inertia I_(zz) about a third (z) axis perpendicular to thefirst axis (x) and the second (y) axis such that I_(XX)<I_(YY) andI_(XX)<I_(ZZ). When the satellite is not in a payload data collectionmode, the satellite is oriented to align the first (x) axis normal tothe plane of the orbit and rotated about the first (x) axis such thatthe second (y) axis is perpendicular to a Sun line of sight. Thesatellite's solar panels are also rotated about the second (y) axis at afirst angle (θ) from the satellite orbital plane to orient the solarpanels towards the Sun. The alignment of the first (x) axis ismaintained normal to the satellite orbital plane during the non-payloaddata collection mode. When the satellite is in a payload data collectionmode, the satellite is rotated about only the first (x) axis to directthe payload at a terrestrial target; and target-related payload data iscollected.

The apparatus comprises a plurality of attitude motive elements, theplurality of attitude motive elements for changing the attitude of thesatellite in the first (x) axis, the second (y) axis, and the third (z)axis and one or more processors, communicatively coupled to the attitudemotive elements. The processor may be communicatively coupled to amemory storing instructions comprising instructions for maneuvering thesatellite to a first orientation by orienting the satellite to align thefirst (x) axis normal to the plane of the orbit, rotating the satelliteabout the first (x) axis such that the second (y) axis is perpendicularto a Sun line of sight, and rotating the solar panels about the second(y) axis at a first angle (θ) from the satellite orbital plane to orientthe solar panels towards the Sun, and maintaining the alignment of thefirst (x) axis normal to the satellite orbital plane during thenon-payload data collection mode when the satellite is not in a payloaddata collection mode. The memory may also store instructions formaneuvering the satellite to a second orientation by rotating thesatellite only about the first (x) axis to direct the payload at aterrestrial target and to collect the payload data when the satellite isin a payload data collection mode.

For the most satellites with global surveillance mission (includingthose in low-earth or mid-earth orbits), the motion of the Sun and theprecession of the satellite's orbital plane are slow compared to theirrelative motion to the ground. Therefore, once the attitude of thesatellite for optimal solar power is set, it will require very little orno adjustment over a small number of orbit revolutions. In typical dailyoperations, the spacecraft will be in the Sun optimal attitude most ofthe time. The attitude may then be changed only to allow optimal payloadsensor coverage when the satellite is near the ground area of interest.

The design described herein places the rotation axis of the solar panelon the y axis of the bus, and the payload sensor array along the x axisof the bus. Due to the length of the array support structure and thecorresponding mass property of the system, the bus x axis is theprincipal axis with minimum moment of inertia. This design places thex-axis perpendicular to the orbital plane to minimize the disturbancetorque due to gravity gradient, and to allow simple rotation about theX-axis for attitude change between optimal Sun and optimal groundcoverage. Small rotations at constant rate about an inertial referencecan be introduced to accommodate the precession of the orbital plane dueto natural perturbation.

BRIEF DESCRIPTION OF THE DRAWINGS

Referring now to the drawings in which like reference numbers representcorresponding parts throughout:

FIG. 1 is a diagram of a representative satellite having a large payloadarray;

FIG. 2 is a block diagram illustrating a satellite attitude controlsystem;

FIG. 3 is a flow chart illustrating representative method steps that canbe used to practice one embodiment of the present invention;

FIG. 4 is a diagram of a satellite in orbit around the Earth in a powercharging configuration, viewed edge-on to its orbital plane;

FIG. 5 is a diagram of the satellite shown in FIG. 4, viewed from alocation perpendicular to the orbital plane;

FIG. 6 is a diagram of a satellite in orbit around the Earth in aservice orientation, viewed edge-on to the orbital plane; and

FIG. 7 is a diagram of the satellite shown in FIG. 6, viewed from alocation perpendicular to the orbital plane.

DETAILED DESCRIPTION OF PREFERRED EMBODIMENTS

In the following description, reference is made to the accompanyingdrawings which form a part hereof, and which is shown, by way ofillustration, several embodiments of the present invention. It isunderstood that other embodiments may be utilized and structural changesmay be made without departing from the scope of the present invention.

FIG. 1 illustrates a three-axis stabilized satellite or spacecraft 100.The spacecraft 100 is preferably situated in a stationary orbit aboutthe Earth. The satellite 100 has a main body 102, a pair of solar panels104, a pair of high gain narrow beam antennas 106, and a telemetry andcommand omnidirectional antenna 108 which is aimed at a control groundstation. The satellite 100 may also include one or more sensors 110 tomeasure the attitude of the satellite 100. These sensors may include sunsensors, earth sensors, and star sensors. Since the solar panels areoften referred to by the designations “North” and “South”, the solarpanels in FIG. 1 are referred to by the numerals 104N and 104S for the“North” and “South” solar panels, respectively. In the illustratedembodiment, the solar panels 104 can be rotated along their longitudinalaxis (the y axis) to direct them at the Sun. Other embodiments mayinclude solar panels that are also rotatable about the z and y axes, butsuch implementations are more expensive, and in light of the inventiondescribed below, unnecessary for purposes of maximizing solar powercollection.

The three axes of the spacecraft 10 are shown in FIG. 1. The pitch axisP lies along the plane of the solar panels 140N and 140S. The roll axisR and yaw axis Y are perpendicular to the pitch axis P and lie in thedirections and planes shown.

The satellite 100 also comprises a payload such as a sensor systemhaving a sensor array 112. In the illustrated embodiment, the sensorarray 112 implements a space-based radar system used to view targets onthe Earth. The sensor array 112 could be coupled to the satellite bus102 via one or more joints and motors that would permit the sensor array112 to be rotated along it's longitudinal (x) axis to view the Earth, ormay be fixedly coupled to the satellite bus 102 so that anyreorientation of the sensor array 112 requires that the satellite bus bereoriented as well. In the illustrated embodiment, the z axis isperpendicular to the sensor array.

Accordingly, in the illustrated embodiment, the satellite has a sensorarray 112 disposed longitudinally along the x axis and rotatable solarpanels 104 disposed longitudinally along a second y axis perpendicularto the x axis. The satellite 100 has a moment of inertia I_(XX) aboutthe x axis, a moment of inertia I_(YY) about the y axis and a moment ofinertia I_(ZZ) about the z axis perpendicular to the first axis (x) andthe second (y) axis. Also in the illustrated embodiment, the massdistribution of the satellite 100 is such that the moment of inertiaabout the x axis is less than the moment of inertia of either the y or zaxes (e.g., I_(XX)<I_(YY) and I_(XX)<I_(ZZ)).

Due to the length and corresponding mass properties of the sensor array112 and support structure the x axis is the axis with the minimum momentof inertia. As described below, this invention advantageously places thex axis perpendicular to the orbital plane 402. This results in twoimportant benefits. First, it minimizes disturbance torques due togravity gradients and second, it allows the satellite to change modesfrom solar collection to data collection and back again through a simplerotation about only the x axis.

FIG. 2 is a diagram depicting the functional architecture of arepresentative attitude control system 200. Control of the satellite 100is provided by a computer or spacecraft control processor (SCP) 202. TheSCP 202 performs a number of functions which may include post ejectionsequencing, transfer orbit processing, acquisition control,stationkeeping control, normal mode control, mechanisms control, faultprotection, and spacecraft systems support, among others. The postejection sequencing could include initializing to assent mode andthruster active nutation control (TANC). The transfer orbit processingcould include attitude data processing, thruster pulse firing, perigeeassist maneuvers, and liquid apogee motor (LAM) thruster firing. Theacquisition control could include idle mode sequencing, sunsearch/acquisition, and Earth search/acquisition. The stationkeepingcontrol could include auto mode sequencing, gyro calibration,stationkeeping attitude control and transition to normal. The normalmode control could include attitude estimation, attitude and solar arraysteering, momentum bias control, magnetic torquing, and thrustermomentum dumping (H-dumping). The mechanisms mode control could includesolar panel control and reflector positioning control. The spacecraftcontrol systems support could include tracking and command processing,battery charge management and pressure transducer processing.

Input to the spacecraft control processor 202 may come from anycombination of a number of spacecraft components and subsystems, such asa transfer orbit sun sensor 204, an acquisition sun sensor 206, aninertial reference unit 208, a transfer orbit Earth sensor 210, anoperational orbit Earth sensor 212, a normal mode wide angle sun sensor214, a magnetometer 216, and one or more star sensors 218.

The SCP 202 generates control signal commands 220 which are directed toa command decoder unit 222. The command decoder unit operates the loadshedding and battery charging systems 224. The command decoder unit alsosends signals to the magnetic torque control unit (MTCU) 226 and thetorque coil 228.

The attitude control system comprises a plurality of attitude motiveelements or actuators that are used to change the attitude of thesatellite in the first (x) axis, the second (y) axis, and the third (z)axis. Such elements include ACS thrusters 236 and momentum wheels 242and 244 of a number and orientation to permit the satellite 100 to berotated about any one or all of the x, y, and z axes. Other attitudecontrol actuators may also be used.

The SCP 202 also sends control commands 230 to the thruster valve driverunit 232 which in turn controls the liquid apogee motor (LAM) thrusters234 and the attitude control thrusters 236.

Wheel torque commands 262 are generated by the SCP 202 and arecommunicated to the wheel speed electronics 238 and 240. These effectchanges in the wheel speeds for wheels in momentum wheel assemblies 242and 244, respectively. The speed of the wheels is also measured and fedback to the SCP 202 by feedback control signal 264.

The spacecraft control processor also sends jackscrew drive signals 266to the momentum wheel assemblies 243 and 244. These signals control theoperation of the jackscrews individually and thus the amount of tilt ofthe momentum wheels. The position of the jackscrews is then fed backthrough command signal 268 to the spacecraft control processor. Thesignals 268 are also sent to the telemetry encoder unit 258 and in turnto the ground station 260.

The spacecraft control processor also sends command signals 254 to thetelemetry encoder unit 258 which in turn sends feedback signals 256 tothe SCP 202. This feedback loop, as with the other feedback loops to theSCP 202 described earlier, assist in the overall control of thespacecraft. The SCP 202 communicates with the telemetry encoder unit258, which receives the signals from various spacecraft components andsubsystems indicating current operating conditions, and then relays themto the ground station 260.

The wheel drive electronics 238, 240 receive signals from the SCP 202and control the rotational speed of the momentum wheels. The jackscrewdrive signals 266 adjust the orientation of the angular momentum vectorof the momentum wheels. This accommodates varying degrees of attitudesteering agility and accommodates movement of the spacecraft asrequired.

The use of reaction wheels or equivalent internal torquers to control amomentum bias stabilized spacecraft allows inversion about yaw of theattitude at will without change to the attitude control. In this sense,the canting of the momentum wheel is entirely equivalent to the use ofreaction wheels.

Other spacecraft employing external torquers, chemical or electricthrusters, magnetic torquers, solar pressure, etc. cannot be invertedwithout changing the control or reversing the wheel spin direction. Thisincludes momentum bias spacecraft that attempt to maintain thespacecraft body fixed and steer payload elements with payload gimbals.

The SCP 202 may include or have access to memory 270, such as a randomaccess memory (RAM). Generally, the SCP 202 operates under control of anoperating system 272 stored in the memory 270, and interfaces with theother system components to accept inputs and generate outputs, includingcommands. Applications running in the SCP 202 access and manipulate datastored in the memory 270. The spacecraft 10 may also comprise anexternal communication device such as a satellite link for communicatingwith other computers at, for example, a ground station. If necessary,operation instructions for new applications can be uploaded from groundstations.

In one embodiment, instructions implementing the operating system 272,application programs, and other modules are tangibly embodied in acomputer-readable medium, e.g., data storage device, which could includea RAM, EEPROM, or other memory device. Further, the operating system 272and the computer program are comprised of instructions which, when readand executed by the SCP 202, causes the spacecraft processor 202 toperform the steps necessary to implement and/or use the presentinvention. Computer program and/or operating instructions may also betangibly embodied in memory 270 and/or data communications devices (e.g.other devices in the spacecraft 10 or on the ground), thereby making acomputer program product or article of manufacture according to theinvention. As such, the terms “program storage device,” “article ofmanufacture” and “computer program product” as used herein are intendedto encompass a computer program accessible from any computer readabledevice or media.

FIG. 3 is a diagram presenting illustrative steps that can be used topractice one embodiment of the present invention. The steps describedtwo modes of operation, each associated with an orientation. In a first(solar collection) mode of operation, the satellite 100 is in a solarenergy collection orientation, and in the second (data collection) modeof operation, the satellite 100 is in a service orientation wherein thesensor array 112 is directed to obtain data from terrestrial objects.

FIG. 3 will be discussed in further reference to FIGS. 4-7. FIGS. 4-5illustrate the satellite 100 in the solar collection mode, while FIGS.6-7 illustrate the satellite 100 in the data collection mode.

Returning now to FIG. 3, block 302 determines whether the satellite 100is in the sensor data collection mode. If the satellite 100 is in thedata processing mode, logic is routed to block 312, where operationsrelated to sensor data collection mode operations are described.

If the satellite 100 is not in the sensor data collection mode, logic isrouted to block 304, where operations related to solar collection aredescribed. If the satellite 100 is not in the data collection mode, thesatellite 100 is oriented to align the x axis normal to the plane of thesatellite's orbit 402, as shown in block 304 and in FIG. 4. In oneembodiment, this is accomplished by orienting the satellite 100 to alignthe x axis to be perpendicular to the Earth nadir (N) 502 and thesatellite velocity vector (V) 504.

To direct the solar panels in the direction of the Sun 406, thesatellite 100 is rotated about the x axis by an angle ψ so that the yaxis is perpendicular to the line of sight to the Sun 403, as shown inblock 306 and FIG. 5. In the illustrated embodiment, this isaccomplished by rotating the satellite bus 102 about the x axis, butthis may be accomplished by rotating the solar panels 104 as well. Thesolar panels 104 may also rotated about the y axis by an angle θ fromthe satellite orbital plane 402, as shown in block 308 and FIG. 4 tofurther orient the solar panels to be perpendicular to the line of sightto the Sun 406. This can accomplished, for example by using motors oractuators in the satellite bus 102 that rotate the solar panels 104about the y axis relative to the satellite bus 102.

While in the solar collection mode, the satellite 100 orientation in thesolar collection mode is maintained by orienting the satellite tomaintain the alignment of the x axis to be normal to the satelliteorbital plane, as shown in block 310. This may be accomplished byrotating the x axis of satellite 100 about the same axis and at the samerate as nodal precession of the orbital plane 402. The satellite 100 canalso be rotated about the x axis to maintain the surface of the solarpanels 104 perpendicular to the Sun line of sight 403. The requiredrotation rate is approximately the reciprocal of the orbital period ofthe satellite 100.

If the satellite 100 is in the sensor data collection or service mode,the satellite 100 is rotated only about the first (x) axis to direct thesensitive axis of the sensor array 112 at one or more terrestrialtargets 602, as shown in block 312 and illustrated in FIGS. 6 and 7.Sensor data is then collected, as illustrated in block 314. After doingso, the surface of the solar panels 104 will no longer be perpendicularto the Sun line of sight 403, so the amount of solar energy collected bythe solar panels will decrease. However, after block 302 determines therequired sensor data is collected, the satellite 100 returns to thesolar collection mode by following steps 304-310 above.

CONCLUSION

This concludes the description of the preferred embodiments of thepresent invention. The foregoing description of the preferred embodimentof the invention has been presented for the purposes of illustration anddescription. It is not intended to be exhaustive or to limit theinvention to the precise form disclosed. Many modifications andvariations are possible in light of the above teaching. For example,while the foregoing was described with respect to a sensor array, theforegoing can be practiced with any payload. Further, although theforegoing is described in terms of being implemented by a processorexecuting instructions stored in a memory, using the foregoing teaching,it is apparent that the invention may also be implemented by one or morespecial purpose processors performing subsets of the defined operations,or by hardware dedicated to the tasks.

It is intended that the scope of the invention be limited not by thisdetailed description, but rather by the claims appended hereto. Theabove specification, examples and data provide a complete description ofthe manufacture and use of the composition of the invention. Since manyembodiments of the invention can be made without departing from thespirit and scope of the invention, the invention resides in the claimshereinafter appended.

1. A method of maneuvering a satellite in an orbit, the satellite having a payload disposed longitudinally along a first (x) axis and rotatable solar panels disposed longitudinally along a second (y) axis perpendicular to the first axis (x), the satellite having a moment of inertia about the first axis I_(XX) and about the second axis I_(YY) and a moment of inertia I_(zz) about a third (z) axis perpendicular to the first axis (x) and the second (y) axis such that I_(XX)<I_(YY) and I_(XX)<I_(ZZ) the method comprising the steps of: when the satellite is not in a payload data collection mode, performing steps comprising the steps of: orienting the satellite to align the first (x) axis normal to the plane of the orbit; rotating the satellite about the first (x) axis such that the second (y) axis is perpendicular to a Sun line of sight; rotating the solar panels about the second (y) axis at a first angle (θ) from the satellite orbital plane to orient the solar panels towards the Sun; and maintaining the alignment of the first (x) axis normal to the satellite orbital plane during the non-payload data collection mode; when the satellite is in a payload data collection mode, performing steps comprising the steps of: rotating the satellite about only the first (x) axis to direct the payload at a terrestrial target; and collecting payload data.
 2. The method of claim 1, wherein the step of orienting the satellite to align the first (x) axis normal to the plane of the orbit comprise instructions for orienting the satellite to align the first axis (x) to be perpendicular to the Earth nadir and a satellite velocity vector.
 3. The method of claim 1, wherein the step of maintaining the alignment of the first (x) axis normal to the satellite orbital plane during the non-payload data collection mode comprises the step of rotating the satellite about the first (x) axis to maintain the alignment of the first (x) axis normal to the satellite orbital plane and to direct the second (y) axis perpendicular to the Sun line of sight.
 4. The method of claim 1, wherein the step of maintaining the alignment of the first (x) axis normal to the satellite orbital plane during the non-payload data collection mode comprises the step of rotating the first (x) axis about the same axis and at the same rate as a nodal precession of the orbital plane.
 5. The method of claim 1, wherein I_(XX)<<I_(ZZ) and I_(XX)<<I_(YY).
 6. The method of claim 1, wherein the payload comprises a sensor array longitudinally disposed along the first (x) axis.
 7. An apparatus for maneuvering a satellite in an orbit, the satellite having a payload disposed longitudinally along a first (x) axis and rotatable solar panels disposed longitudinally along a second (y) axis perpendicular to the first axis (x), the satellite having a moment of inertia about the first axis I_(XX) and about the second axis I_(YY) and a moment of inertia I_(zz) about a third axis (z) perpendicular to the first axis (x) and the second (y) axis such that I_(XX)<I_(YY) and I_(XX)<I_(ZZ), comprising an attitude control system, comprising a plurality of attitude motive elements, the plurality of attitude motive elements for changing the attitude of the satellite in the first (x) axis, the second (y) axis, and the third (z) axis; a processor, communicatively coupled to the attitude motive elements, the processor also communicatively coupled to a memory storing instructions comprising instructions for: when the satellite is not in a payload data collection mode, maneuvering the satellite to a first orientation by orienting the satellite to align the first (x) axis normal to the plane of the orbit, rotating the satellite about the first (x) axis such that the second (y) axis is perpendicular to a Sun line of sight, and rotating the solar panels about the second (y) axis at a first angle (θ) from the satellite orbital plane to orient the solar panels towards the Sun, and maintaining the alignment of the first (x) axis normal to the satellite orbital plane during the non-payload data collection mode; and when the satellite is in a payload data collection mode, maneuvering the satellite to a second orientation by rotating the satellite only about the first (x) axis to direct the payload at a terrestrial target to collect the payload data.
 8. The apparatus of claim 7, wherein the instructions for orienting the satellite to align the first (x) axis normal to the plane of the orbit comprise instructions for orienting the satellite to align the first axis (x) to be perpendicular to the Earth nadir and a satellite velocity vector.
 9. The apparatus of claim 7, wherein the instructions for maintaining the alignment of the first (x) axis normal to the satellite orbital plane during the non-payload data collection mode comprises instructions for rotating the satellite about the first (x) axis to maintain the alignment of the first (x) axis normal to the satellite orbital plane and to direct the second (y) axis perpendicular to the Sun line of sight.
 10. The apparatus of claim 7, wherein the instructions for maintaining the alignment of the first (x) axis normal to the satellite orbital plane during the non-payload data collection mode comprise instructions for rotating the first (x) axis about the same axis and at the same rate as a nodal precession of the orbital plane.
 11. The apparatus of claim 7, wherein I_(XX)<<I_(ZZ) and I_(XX)<<I_(YY).
 12. The apparatus of claim 7, wherein the payload comprises a sensor array longitudinally disposed along the first (x) axis.
 13. A apparatus for maneuvering a satellite in an orbit, the satellite having a payload disposed longitudinally along a first (x) axis and rotatable solar panels disposed longitudinally along a second (y) axis perpendicular to the first axis (x), the satellite having a moment of inertia about the first axis I_(XX) and about the second axis I_(YY) and a moment of inertia I_(zz) about a third axis (z) perpendicular to the first axis (x) and the second (y) axis such that I_(XX)<I_(YY) and I_(XX)<I_(ZZ) the method comprising: means for orienting the satellite to align the first (x) axis normal to the plane of the orbit, maintaining the alignment of the first (x) axis normal to the satellite orbital plane during the non-payload data collection mode, and rotating the satellite about the first (x) axis such that the second (y) axis is perpendicular to a Sun line of sight, and rotating the solar panels about the second (y) axis at a first angle (θ) from the satellite orbital plane to orient the solar panels towards the Sun when the satellite is not in a payload data collection mode, maneuvering the satellite to a first orientation; and means for rotating the satellite only about the first (x) axis to direct the payload at a terrestrial target and collecting payload data when the satellite is in a payload data collection mode, maneuvering the satellite to a second orientation.
 14. The apparatus of claim 13, wherein the means for orienting the satellite to align the first (x) axis normal to the plane of the orbit comprise means for orienting the satellite to align the first axis (x) to be perpendicular to the Earth nadir and a satellite velocity vector.
 15. The method of claim 1, wherein the means for maintaining the alignment of the first (x) axis normal to the satellite orbital plane during the non-payload data collection mode comprises means for rotating the satellite about the first (x) axis to maintain the alignment of the first (x) axis normal to the satellite orbital plane and to direct the second (y) axis perpendicular to the Sun line of sight.
 16. The method of claim 1, wherein the means for maintaining the alignment of the first (x) axis normal to the satellite orbital plane during the non-payload data collection mode comprises the means for rotating the first (x) axis about the same axis and at the same rate as a nodal precession of the orbital plane.
 17. The method of claim 1, wherein I_(XX)<<I_(ZZ) and I_(XX)<<I_(YY).
 18. The method of claim 1, wherein the satellite comprises a payload array longitudinally disposed along the first (x) axis. 